A gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section and an exhaust section. In operation, air enters an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section through a hot gas path defined within the turbine section and then exhausted from the turbine section via the exhaust section.
In particular configurations, the turbine section includes, in serial flow order, a high pressure (HP) turbine and a low pressure (LP) turbine. The HP turbine and the LP turbine each include various rotatable turbine components such as turbine rotor disks and turbine rotor blades, and various stationary turbine components such as stator vanes or nozzles, turbine shrouds, shroud supports and engine frames. The rotatable and the stationary turbine components at least partially define the hot gas path through the turbine section. As the combustion gases flow through the hot gas path, thermal energy is transferred from the combustion gases to the rotatable turbine components and the stationary turbine components. As a result, it is generally necessary to cool the various rotatable and stationary turbine components to meet thermal and/or mechanical performance requirements.
Conventionally, a cooling medium such as compressed air is routed from the compressor section through various cooling passages or circuits defined within or around the various rotatable and stationary turbine components, thus providing cooling to those components. However, use of a 360 degree ceramic matrix composite shroud in the HPT allows for less backside shroud cooling via the compressed air due to the favorable thermal properties of the ceramic matrix composite material. As a result, the amount of compressed air normally routed into or around the shroud assembly is reduced, thus enhancing overall engine performance and/or efficiency.
The reduction in cooling flow to the shroud assembly may result in higher temperatures on stationary hardware such as the shroud support hardware of the shroud assembly that is potentially exposed to or in the line of sight of the combustion gases flowing through the hot gas path. Increased thermal stresses on the shroud support hardware generally occurs when the shroud support hardware is formed from metal or other materials having less favorable thermal properties for exposure to the combustion gases than the ceramic matrix composite material used for the shroud. Accordingly, a turbine shroud assembly configured to thermally shield the shroud support and/or other adjacent stationary hardware to reduce thermal stresses would be welcomed in the technology.